Method and system for detecting forces on aircraft

ABSTRACT

A method for sensing a force applied to an aircraft includes receiving a derivative of the acceleration of a motion of a portion of the aircraft, determining whether the derivative of the acceleration of the motion of the portion of the aircraft exceeds a threshold, and outputting an indication that a force has been applied to the portion of the aircraft responsive to determining that the derivative of the acceleration of motion of the portion of the aircraft exceeds the threshold.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. Provisional Application No.61/255,547, filed Oct. 28, 2009.

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to detecting impact forceson aircraft, and in particular to detecting landing gear impact onaircraft.

Aircraft such as, for example, rotary wing aircraft and fixed wingaircraft use a variety of sensors to provide feedback to aircraftcontrol systems. Detecting when a force, such as weight, is applied tothe landing assemblies or other portions of an aircraft provides usefulfeedback to aircraft systems. Previous systems used sensors located oneach landing assembly to determine whether weight was applied to alanding assembly. The use of these sensors increased the weight andcomplexity of the aircraft, and had limited fidelity in sensing actualweight applied to a landing assembly.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the invention, a method for sensing a forceapplied to an aircraft includes receiving a derivative of theacceleration of a motion of a portion of the aircraft, determiningwhether the derivative of the acceleration of the motion of the portionof the aircraft exceeds a threshold, and outputting an indication that aforce has been applied to the portion of the aircraft responsive todetermining that the derivative of the acceleration of motion of theportion of the aircraft exceeds the threshold.

According to another aspect of the invention, a method for sensing atakeoff of an aircraft includes receiving a rate of change in thevertical motion of the aircraft, determining whether the rate of changein the vertical motion of the aircraft exceeds a first threshold,integrating the rate of change in the vertical motion of the aircraftand outputting a virtual altitude signal, responsive to receiving theindication that the portion of the aircraft is contacting a surface,delaying the virtual altitude signal through a discrete low pass filterand outputting the delayed virtual altitude signal, subtracting thedelayed virtual altitude signal from the virtual altitude signal tooutput an altitude perturbation signal, determining whether the altitudeperturbation signal exceeds a second threshold value, and outputting anindication that the portion of the aircraft is not contacting thesurface responsive to determining that the rate of change in thevertical motion of the aircraft exceeds the first threshold anddetermining that the altitude perturbation signal exceeds the secondthreshold value.

According to yet another aspect of the invention, a system for sensing aforce applied to an aircraft includes a sensor, and a processoroperative to receive a signal indicative of an acceleration of a motionof the aircraft, apply a kinematic equation to the first signal totransform the indication of the acceleration of the motion of theaircraft to indicate an acceleration of a motion of a portion of theaircraft, calculate a derivative of the acceleration of the motion ofthe portion of the aircraft, determine whether the derivative of theacceleration of the motion of the portion of the aircraft exceeds athreshold, output an indication that a force has been applied to theportion of the aircraft responsive to determining that the derivative ofthe acceleration of the motion of the portion of the aircraft exceedsthe threshold.

These and other advantages and features will become more apparent fromthe following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWING

The subject matter which is regarded as the invention is particularlypointed out and distinctly claimed in the claims at the conclusion ofthe specification. The foregoing and other features, and advantages ofthe invention are apparent from the following detailed description takenin conjunction with the accompanying drawings in which:

FIG. 1 illustrates a block diagram of an exemplary embodiment of anaircraft 100.

FIG. 2 illustrates a block diagram of an exemplary embodiment of logicperformed by the processor of FIG. 1.

FIG. 3 illustrates an example of the geometric relationship between asensor and a nose landing assembly of FIG. 1.

FIG. 4 illustrates a block diagram of exemplary impact detection logicof FIG. 2.

FIG. 5 illustrates a block diagram of exemplary takeoff detection logicof FIG. 2.

The detailed description explains embodiments of the invention, togetherwith advantages and features, by way of example with reference to thedrawings.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 illustrates a block diagram of an exemplary embodiment of anaircraft 100. The aircraft 100 includes a nose landing assembly 101, aleft landing assembly 103, and a right landing assembly 105. The landingassemblies may include, for example, a landing gear assembly thatincludes an inflatable wheel, or any other device that is operative tocontact a landing surface. For example a skid assembly may be used, andportions of the skid assembly may be designated as contact pointssimilar to the gear described above. The aircraft 100 includes aprocessor 102 that is communicatively connected to flight controls 104and sensors 106 that may include, for example, a gyro sensor, one ormore accelerometers, a global positioning system (GPS), or any otherinertial sensors. The processor 102 may also be communicativelyconnected to a memory 110 and a display 108.

FIG. 2 illustrates a block diagram of an exemplary embodiment of logicperformed by the processor 102. In this regard, the processor 102receives input data from the sensors 106. The input includesacceleration (ax, ay, and az) from, for example, an accelerometer,velocity (U, V, W) from, for example, a GPS or derived from anaccelerometer, orientation (pitch, roll, yaw; θ, σ, φ) from for example,a gyroscope (gyro), and a rate of change in orientation (P, Q, R) from,for example, a gyro. In block 202 the signals are processed tomathematically transform vectors associated with the signals at thelocation of the sensors to positions associated with each gear. Forexample, the accelerometer may be located close to the center of mass ofthe aircraft 100, however the gear are located geometrically indifferent locations. The geometric relationship between theaccelerometer and a particular gear may be measured or known, allowingthe input from the accelerometer to be mathematically transformed usinga kinematic relationship such that the transformed inputs representacceleration at a particular gear. The processed sensor data is sent toimpact detection logic 204, for landing evolutions, or takeoff detectionlogic 206, for takeoff evolutions. The impact detection logic 204 andtakeoff detection logic 206 output a signal to the force on gear logic208 that outputs a force on gear signal 210. The force on gear signal210 indicates that a weight on wheel force has been applied to a gear.The indication provides information to the aircraft 100 operator and/orautomatic control systems of the aircraft 100 that assists in operatingthe aircraft. Particularly, the weight on wheel force may indicate thatthe aircraft has landed or has taken off from a landing area.

FIG. 3 illustrates an example of the geometric relationship between asensor 106 and the nose landing assembly 101 including an example ofcoordinate systems that are associated with the sensor 106 and the noselanding assembly 101. A kinematic transform may be used tomathematically associate the data collected by the sensor 106 to thenose landing assembly 101. Thus, for example, a movement sensed by thesensor 106 in the X₁ direction, may be kinematically transformed to anassociate the movement with a force applied to the nose landing assembly101. A vector representing the force applied to the nose landingassembly 101 may be plotted on the X₂, Y₂, Z₂ coordinate system.

FIG. 4 illustrates a block diagram of exemplary impact detection logic204 (of FIG. 2) used to determine if a force has been applied to a gearon the aircraft 100. The logic 204 may be applied in a similar manner toeach gear. For exemplary purposes, the description below will describelogic used to determine whether a force or weight has been applied tothe nose landing assembly 101 (of FIG. 1), however the logic may beapplied simultaneously to any landing assembly or portion of a landingassembly. In this regard, if the aircraft landing is expected, a signal408 is output that cues the impact detection logic 204. Vertical jerkdata 402 is compared to a vertical jerk threshold value 401. Thevertical jerk data is the derivative of the acceleration in a verticaldirection. If the vertical jerk data 402 is greater than the verticaljerk threshold value 401 a signal indicating that the threshold isexceeded is output. Rolling jerk data 404 is compared to a rolling jerkthreshold value 403. Rolling jerk data 404 is a derivative of theacceleration of the roll. If the rolling jerk data 404 is greater thanthe rolling jerk threshold value 403, a signal indicating that thethreshold is exceeded is output. Pitching jerk data 406 is compared to apitching jerk threshold 405. The pitching jerk data 406 is a derivativeof the acceleration of the pitch. If the pitching jerk data 406 isgreater than the pitching jerk threshold value 405, a signal indicatingthat the threshold is exceeded is output. The signals are output to anAND logic that determines whether each of the three thresholds have beenexceeded. The force on gear logic 208 outputs a force on gear signal210, set to true, that indicates that a force has been applied to thenose landing assembly 101. If a takeoff signal 410 is output (true) bythe takeoff detection logic 206, it resets the force on gear signal 210to false.

The illustrated embodiment above describes the logic associated with thenose landing assembly 101, however the logic may be used to determine animpact, force, or weight that is applied to any gear, or location on theaircraft 100. Regarding the nose landing assembly 101, a force from theground (or weight) creates positive pitching signals and negativevertical jerk signals. A force (or weight) on the left gear 103 createsa positive rolling jerk signal and negative pitching jerk and negativevertical jerk signals. A force (or weight) on the right gear 105 createsnegative rolling jerk, negative pitching jerk, and negative verticaljerk signals. The thresholds may be determined by design parameters, andthe geometry of the aircraft 100.

FIG. 5 illustrates a block diagram of exemplary embodiment of takeoffdetection logic 206 (of FIG. 2). A vertical rate of acceleration of thegear 502 is compared with a vertical rate threshold 501. A signal 503 isoutput if the vertical rate of acceleration of the gear 502 is greaterthan the vertical rate threshold 501. The forces on the gear signals 508are compared using OR logic, if either gear force signal is true,indicating ground contact condition, the vertical rate of the gear 502is integrated to output a virtual altitude signal 510. The signal 510 isdelayed through a discrete low pass filter in block 506 outputting adelayed virtual altitude signal 511. Altitude perturbation signal 512 iscomputed by subtracting the delayed virtual altitude signal 511 from thevirtual altitude signal 510 and then is compared to an altitudethreshold 505. If the value of the altitude perturbation signal 512 isgreater than the altitude threshold 505, a signal 507 is output. If thesignals 503 and 507 are received at the force on ground logic 208,takeoff signals 509 and 410 are output (set to true). If neither gearforce signal is true, indicating in air condition, the integrator inputis set to zero thus disabling the take off detection.

While the invention has been described in detail in connection with onlya limited number of embodiments, it should be readily understood thatthe invention is not limited to such disclosed embodiments. Rather, theinvention can be modified to incorporate any number of variations,alterations, substitutions or equivalent arrangements not heretoforedescribed, but which are commensurate with the spirit and scope of theinvention. Additionally, while various embodiments of the invention havebeen described, it is to be understood that aspects of the invention mayinclude only some of the described embodiments. Accordingly, theinvention is not to be seen as limited by the foregoing description, butis only limited by the scope of the appended claims.

What is claimed is:
 1. A method for sensing a force applied to anaircraft comprising: receiving a derivative of the acceleration of amotion of a portion of the aircraft; determining, by a processor,whether the derivative of the acceleration of the motion of the portionof the aircraft exceeds a threshold; and outputting, by the processor,an indication that a force has been applied to the portion of theaircraft responsive to determining that the derivative of theacceleration of motion of the portion of the aircraft exceeds thethreshold, wherein the force applied to the portion of the aircraft is aweight on wheel force indicating landing of the aircraft; the processorresetting the indication that the force has been applied to the portionof the aircraft upon receiving a takeoff signal, the takeoff signalindicating the aircraft has taken flight.
 2. The method of claim 1,wherein the derivative of the acceleration of the motion of the portionof the aircraft is the derivative of the acceleration of a verticalmotion of the portion of the aircraft and the threshold is associatedwith the vertical motion of the portion of the aircraft.
 3. The methodof claim 2, wherein the derivative of the acceleration of the verticalmotion of the portion of the aircraft is calculated by: receiving afirst signal indicative of the acceleration of the vertical motion ofthe aircraft from a sensor; applying a kinematic equation to the firstsignal to transform the indication of the acceleration of the verticalmotion of the aircraft to indicate the acceleration of the verticalmotion of the portion of the aircraft; and taking a derivative of theacceleration of the vertical motion of the portion of the aircraft. 4.The method of claim 2, wherein the threshold is a negative value.
 5. Themethod of claim 1, wherein the derivative of the acceleration of themotion of the portion of the aircraft is the derivative of theacceleration of a rolling motion of the portion of the aircraft and thethreshold is associated with the rolling motion of the portion of theaircraft.
 6. The method of claim 5, wherein the derivative of theacceleration of the rolling motion of the portion of the aircraft iscalculated by: receiving a second signal indicative of the accelerationof the rolling motion of the aircraft from a sensor; applying akinematic equation to the second signal to transform the indication ofthe acceleration of the rolling motion of the aircraft to indicate theacceleration of the rolling motion of the portion of the aircraft; andtaking a derivative of the acceleration of the rolling motion of theportion of the aircraft.
 7. The method of claim 5, wherein the portionof the aircraft is a left landing assembly and the threshold is apositive value.
 8. The method of claim 5, wherein the portion of theaircraft is a right landing assembly and the threshold is a negativevalue.
 9. The method of claim 1, wherein the derivative of theacceleration of the motion of the portion of the aircraft is thederivative of the acceleration of a pitching motion of the portion ofthe aircraft and the threshold is associated with the pitching motion ofthe portion of the aircraft.
 10. The method of claim 9, wherein thederivative of the acceleration of the pitching motion of the portion ofthe aircraft is calculated by: receiving a third signal indicative ofthe acceleration of the pitching motion of the aircraft from a sensor;applying a kinematic equation to the third signal to transform theindication of the acceleration of the pitching motion of the aircraft toindicate the acceleration of the pitching motion of the portion of theaircraft; and taking a derivative of the acceleration of the pitchingmotion of the portion of the aircraft.
 11. The method of claim 9,wherein the portion of the aircraft is a nose landing assembly and thethreshold is a positive value.
 12. The method of claim 9, wherein thethreshold is a negative value.
 13. A system for sensing a force appliedto an aircraft comprising: a sensor; and a processor operative toreceive a signal indicative of an acceleration of a motion of theaircraft, apply a kinematic equation to the first signal to transformthe indication of the acceleration of the motion of the aircraft toindicate an acceleration of a motion of a portion of the aircraft,calculate a derivative of the acceleration of the motion of the portionof the aircraft, determine whether the derivative of the acceleration ofthe motion of the portion of the aircraft exceeds a threshold, output anindication that a force has been applied to the portion of the aircraftresponsive to determining that the derivative of the acceleration of themotion of the portion of the aircraft exceeds the threshold, wherein theforce applied to the portion of the aircraft is a weight on wheel forceindicating landing of the aircraft; the processor resetting theindication that the force has been applied to the portion of theaircraft upon receiving a takeoff signal, the takeoff signal indicatingthe aircraft has taken flight.
 14. The system of claim 13, wherein thederivative of the acceleration of the motion of the portion of theaircraft is the derivative of the acceleration of a vertical motion ofthe portion of the aircraft and the threshold is associated with thevertical motion of the portion of the aircraft.
 15. The system of claim13, wherein the derivative of the acceleration of the motion of theportion of the aircraft is the derivative of the acceleration of arolling motion of the portion of the aircraft and the threshold isassociated with the rolling motion of the portion of the aircraft. 16.The system of claim 13, wherein the derivative of the acceleration ofthe motion of the portion of the aircraft is the derivative of theacceleration of a pitching motion of the portion of the aircraft and thethreshold is associated with the pitching motion of the portion of theaircraft.
 17. The system of claim 13, wherein the portion of theaircraft is a landing assembly.
 18. A system for sensing a force appliedto an aircraft comprising: a sensor; and a processor operative toreceive a signal indicative of an acceleration of a motion of theaircraft, apply a kinematic equation to the first signal to transformthe indication of the acceleration of the motion of the aircraft toindicate an acceleration of a motion of a portion of the aircraft,calculate a derivative of the acceleration of the motion of the portionof the aircraft, determine whether the derivative of the acceleration ofthe motion of the portion of the aircraft exceeds a threshold, output anindication that a force has been applied to the portion of the aircraftresponsive to determining that the derivative of the acceleration of themotion of the portion of the aircraft exceeds the threshold; wherein thederivative of the acceleration of the motion of the portion of theaircraft includes the derivative of the acceleration of a verticalmotion of the portion of the aircraft and the threshold is associatedwith the vertical motion of the portion of the aircraft; wherein thederivative of the acceleration of the motion of the portion of theaircraft includes the derivative of the acceleration of a rolling motionof the portion of the aircraft and the threshold is associated with therolling motion of the portion of the aircraft; wherein the derivative ofthe acceleration of the motion of the portion of the aircraft includesthe derivative of the acceleration of a pitching motion of the portionof the aircraft and the threshold is associated with the pitching motionof the portion of the aircraft.